Propulsion system using large scale vortex generators for flow redistribution and supersonic aircraft equipped with the propulsion system

ABSTRACT

An arrangement for use with a propulsion system for a supersonic aircraft includes a center body configured for coupling to an inlet and to support a boundary layer formed when the supersonic aircraft is flown at a predetermined altitude supersonic speed. The arrangement further includes a first vortex generator disposed on the center body. The first vortex generator extends a first height above the center body. The arrangement still further includes a second vortex generator disposed on the center body. The second vortex generator extends a second height above the center body, the second height being greater than the first height. The first height and the second height are greater than approximately seventy-five percent of a thickness of the boundary layer proximate a location of the first vortex generator and the second vortex generator, respectively, when the aircraft if flown at the predetermined altitude and the predetermined speed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No.14/172,560 filed Feb. 4, 2014, and entitled “Propulsion System UsingLarge Scale Vortex Generators for Flow Redistribution and SupersonicAircraft Equipped with the Propulsion System” which claims the benefitof U.S. Provisional Application No. 61/764,658 filed Feb. 14, 2013, andentitled “Large-Scale Flow Redistribution Using Vortex Generators”, bothof which are incorporated herein by reference in their entirety.

TECHNICAL FIELD

The present invention generally relates to aviation and moreparticularly relates to a propulsion system for a supersonic aircraft.

BACKGROUND

The engine of a conventional supersonic aircraft includes a center bodyhaving an elongated compression surface to help improve pressurerecovery caused by movement of the engine through the air at supersonicspeeds. The compression surface, together with other features of theaircraft's propulsion system, slows the supersonic airflow entering thepropulsion system to a speed that is compatible with the turbo machineryof the engine.

One undesirable consequence of having an elongated compression surfaceis the buildup of a relatively thick boundary layer on internal surfacesof the inlet (e.g., portions of the diffuser). The boundary layer is aportion of the airflow located proximate a viscous surface (such as thecompression surface and the surface of the diffuser) that, because ofits interaction with the viscous body, moves slower than the free streamvelocity.

Because the boundary layer air is moving at a slower speed than theremainder of the airflow, the boundary layer air will have a lowerstagnation pressure than the remainder of the airflow. This leads todistortion in the stagnation pressure of the airflow entering theAerodynamic Interaction Plane (“AIP”) (e.g., the fan or the face of theengine). This distortion in stagnation pressure is undesirable, becauseit may adversely impact both engine operability and performance.

Several different solutions have been developed to combat the distortionin the stagnation pressure caused by the elongated compression surface.For example, some propulsion systems bleed the boundary layer from theairflow by passing the airflow over a porous surface and using lowpressure to extract the boundary layer from the airflow. While this iseffective at diminishing the thickness of the boundary layer, such bleedsystems add cost, complexity, and weight to a propulsion system.

Another solution has been to position vortex generators of modest heighton the center body. These vortex generators have a height ranging fromtwenty percent to forty percent of the local boundary layer thicknessand generate vortices that propagate completely within the boundarylayer. These vortices increase the energy level of the boundary layerwhich, in turn, allows the boundary layer to remain more robustlyattached to the curving surface of the center body or other inletsurface. While this inhibits growth and separation of the boundarylayer, it does not modify its structure or appreciably reduce itsthickness and the stagnation pressure of the air entering the AIPremains distorted.

Accordingly, it is desirable to provide a propulsion system that reducesdistortion of the stagnation pressure of the airflow entering the AIP.Additionally, it is desirable to provide a supersonic aircraft equippedwith a propulsion system that reduces the distortion of the stagnationpressure. Furthermore, other desirable features and characteristics willbecome apparent from the subsequent summary and detailed description andthe appended claims, taken in conjunction with the accompanying drawingsand the foregoing technical field and background.

BRIEF SUMMARY

A propulsion system for a supersonic aircraft and an arrangement for usewith a propulsion system for a supersonic aircraft are disclosed herein.

In a first non-limiting embodiment, the arrangement includes, but is notlimited to, a center body configured for coupling to an inlet and tosupport a boundary layer formed when the supersonic aircraft is flown ata predetermined altitude and a predetermined speed greater than Mach 1.The arrangement further includes, but is not limited to, a first vortexgenerator disposed on the center body. The first vortex generatorextends a first height above the center body. The arrangement stillfurther includes, but is not limited to, a second vortex generatordisposed on the center body. The second vortex generator extends asecond height above the center body, the second height being greaterthan the first height. The first height and the second height aregreater than approximately seventy-five percent of a thickness of theboundary layer proximate a location of the first vortex generator andthe second vortex generator, respectively, when the aircraft if flown atthe predetermined altitude and the predetermined speed.

In another non-limiting embodiment, the propulsion system includes, butis not limited to, a plurality of vortex generators positioned upstreamof an engine of the propulsion system. Each vortex generator of theplurality of vortex generators has a height such that when thesupersonic aircraft is flown at a predetermined altitude and at apredetermined cruise speed, the predetermined cruise speed being greaterthan Mach 1, the plurality of vortex generators create a plurality ofvortices that propagate at least partially outside of a boundary layerformed proximate a surface of a supersonic inlet. The plurality ofvortices causes a high-velocity portion of the airflow to move towardsthe engine core and a low-velocity portion of the airflow to movetowards the engine bypass prior to the airflow reaching a face of theengine. The plurality of vortex generators are disposed aft of aterminal shock formed when the supersonic aircraft is flown at thepredetermined cruise speed. Each vortex generator of the plurality ofvortex generators has a height greater than a thickness of the boundarylayer at the location of each vortex generator of the plurality ofvortex generators.

In another non-limiting embodiment, the propulsion system includes, butis not limited to, a plurality of vortex generators positioned upstreamof an engine of the propulsion system. Each vortex generator of theplurality of vortex generators has a height such that when thesupersonic aircraft is flown at a predetermined altitude and at apredetermined cruise speed, the predetermined cruise speed being greaterthan Mach 1, the plurality of vortex generators create a plurality ofvortices that propagate at least partially outside of a boundary layerformed proximate a surface of a supersonic inlet. The plurality ofvortices causes a high-velocity portion of the airflow to move towardsthe surface and a low-velocity portion of the airflow to move away fromthe surface prior to the airflow reaching a face of the engine. Theplurality of vortex generators are disposed aft of a terminal shockformed when the supersonic aircraft is flown at the predetermined cruisespeed. Each vortex generator of the plurality of vortex generators havea height greater than a thickness of the boundary layer at the locationof each vortex generator of the plurality of vortex generators.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention will hereinafter be described in conjunction withthe following drawing figures, wherein like numerals denote likeelements, and

FIG. 1 is a schematic cross-sectional view illustrating a prior artpropulsion system for use with a supersonic aircraft;

FIG. 2 is a schematic cross-sectional view illustrating a non-limitingembodiment of a propulsion system made in accordance with the teachingsof the present disclosure;

FIG. 3 is a cross-sectional view taken across the line 3-3 of FIG. 2;and

FIG. 4 is a fragmentary schematic view illustrating a portion of thepropulsion system of FIG. 2 from the perspective of arrow 4 of FIG. 3;

FIG. 5 is a fragmentary perspective view illustrating a firstnon-limiting arrangement of a plurality of vortex generators includedwith the propulsion system of FIG. 2;

FIG. 6 is a fragmentary perspective view illustrating a secondnon-limiting arrangement of a plurality of vortex generators includedwith the propulsion system of FIG. 2;

FIG. 7 is an expanded schematic side view illustrating a portion of thepropulsion system of FIG. 2;

FIG. 8 is a schematic cross-sectional view illustrating anothernon-limiting embodiment of a propulsion system made in accordance withthe teachings of the present disclosure;

FIG. 9 is a cross-sectional view taken across the line 9-9 of FIG. 8;

FIG. 10 is a schematic cross-sectional view illustrating yet anothernon-limiting embodiment of a propulsion system made in accordance withthe teachings of the present disclosure; and

FIG. 11 is a perspective view illustrating a non-limiting embodiment ofa supersonic aircraft equipped with a propulsion system made inaccordance with the teachings of the present disclosure.

DETAILED DESCRIPTION

The following detailed description is merely exemplary in nature and isnot intended to limit the invention or the application and uses of theinvention. Furthermore, there is no intention to be bound by any theorypresented in the preceding background or the following detaileddescription.

An improved propulsion system for a supersonic aircraft and a supersonicaircraft equipped with the propulsion system are disclosed herein. In anexemplary embodiment, the propulsion system of the present disclosureincludes an engine, an inlet center body having an elongated compressionsurface, a shroud that partially surrounds both the engine and thecenter body and a plurality of vortex generators. The shroud isconfigured to guide an airflow passing over the center body towards theengine where it is then ingested by the engine and may comprise anacelle, a bypass splitter, or any other suitable structure.

The vortex generators of the present disclosure are relatively tall ascompared with conventional vortex generators. While conventional vortexgenerators have a height that is only a small fraction of the thicknessof the boundary layer that forms on the center body during supersonicflight at a predetermined speed (e.g., supersonic cruise), the vortexgenerators of the present disclosure have a height that approaches orexceeds the thickness of the boundary layer at the location where theboundary layer encounters the vortex generators. In an example, thevortex generators of the present disclosure have a height of at leastseventy five percent of the thickness of the boundary layer that formson the center body when the aircraft is flown at the predeterminedspeed. This extended height will allow the vortex generators to reach aregion of the boundary layer where the air has a relatively highvelocity as compared with lower regions of the boundary layer where theair's velocity can be quite low. In other examples, the vortexgenerators will have an even taller profile (e.g., having a height thatis equal to, or greater than the thickness of the boundary layer at thelocation where the boundary layer encounters the vortex generator) andwill therefore encounter portions of the airflow having even highervelocities.

By extending into a region of the airflow where the air is moving at arelatively high velocity, the vortex generators of the presentdisclosure are able to generate vortices that propagate outside of theboundary layer. This is a significant departure from conventional vortexgenerators which, because of their lower profiles, generate vorticesthat propagate entirely within the boundary layer. Conventionally,vortices that propagate solely within the boundary layer are desirablebecause they enable the boundary layer to remain more robustly attachedto the center body and thereby counter the natural tendency of theboundary layer to thicken as it passes over curved surfaces.

In the present disclosure, however, the relatively tall vortexgenerators give rise to vortices that propagate outside of the boundarylayer when the supersonic aircraft is flown at the predetermined speed.When arranged in pairs having opposite angles of incidence, the vortexgenerators are capable of producing local up-wash and downwash as aresult of the interaction of adjacent counter-rotating vortices. Theheight of the vortex generators in the present invention increases thestrength of this effect, due to the higher flow velocities encounteredby the taller vortex generators. The vortex generators of the presentdisclosure are therefore able to move the higher speed air of thenon-boundary layer portion of the airflow radially inward towards thecenter body and also move the lower speed air of the boundary layerradially outward and away from the center body. Whereas the goal ofconventional vortex generators has been to keep the boundary layerattached to the center body, one of the goals of the vortex generatorsof the present disclosure is to cause the boundary layer to detach fromthe center body and to have that slower moving air move up into theremainder of the airflow where it can mix with the higher speed air.This mixing reduces the variation in stagnation pressure of the airflowin the radial direction and thereby presents the AIP (e.g., the engineface) with an airflow having less radial distortion than would otherwiseresult from the use of conventional vortex generators (or from the useof no vortex generators). By arranging the vortex generatorscircumferentially around the center body, or around the inner surface ofthe shroud, or both, the circumferential distortion of the stagnationpressure can also be reduced.

It is desirable to present an engine face with an airflow having ahomogenous stagnation pressure rather than an airflow having radialdistortion in stagnation pressure. Engines have sensitivities todistortion in the flow and by making the stagnation pressure across theentire flow more homogenous, the engine's overall performance may beimproved. In addition, depending on the engine that is utilized in thepropulsion system, there can also be an increase in engine thrust. Forexample, if the engine has an engine cycle wherein the core is moresensitive to changes in stagnation pressure than is the engine's fan andbypass, then by moving the slower moving air of the boundary layer awayfrom the center body and by moving the faster moving air of theremainder of the airflow towards the center body, the core will bepresented with an airflow having a higher velocity and hence a higherstagnation pressure. This, in turn, can enhance the amount of thrust theengine generates.

A greater understanding of the propulsion system described above and ofthe supersonic aircraft employing the propulsion system may be obtainedthrough a review of the illustrations accompanying this applicationtogether with a review of the detailed description that follows.

FIG. 1 is a schematic cross-sectional view illustrating a prior artpropulsion system 20. Prior art propulsion system 20 includes a centerbody 22 having an elongated compression surface 24, a nacelle 26including a cowl 27, an inlet 28 formed by the separation between centerbody 22 and cowl 27, a splitter 30 and an engine 32 having an enginecore 34 and an engine bypass 36.

A supersonic free stream of air 39 approaching propulsion system 20 willinitially encounter elongated compression surface 24. Elongatedcompression surface 24 will cause supersonic free stream of air 39 tochange direction and follow the contours of the compression surface.This interaction decelerates supersonic free stream air 39 and causes aboundary layer 38 to form on center body 22. Boundary layer 38 is aregion of stagnant and slower moving air whose thickness increases thefurther it travels along center body 22.

As supersonic free stream of air 39 reaches inlet 28, it passes througha terminal shock 40 extending from cowl 27 to center body 22. Passagethrough terminal shock 40 decelerates the airflow to subsonic speedswhich are more compatible with engine 32. Further slowing may berequired and if so, will occur downstream of terminal shock 40.

As the airflow continues aft of inlet 28, a leading edge 44 of splitter30 divides the airflow into a primary airflow 45 and a secondary airflow47. Primary airflow 45 is guided or routed by splitter 30 towards engine32 while secondary airflow 47 moves through a bypass channel 42 thatavoids engine 32 altogether.

As primary airflow 45 moves downstream past a leading edge 44 ofsplitter 30, it enters a diffuser region 46. In diffuser region 46,center body 22 narrows, creating a larger volume of space for primaryairflow 45 to move through. This narrowing of center body 22 causesboundary layer 38 to thicken as it attempts to remain entrained to thesurface of center body 22.

When primary airflow 45 arrives at face 48 of engine 32, boundary layer38 is at its thickest condition. The portion of primary airflow 45 thatis outside of boundary layer 38 is moving at a high velocity andtherefore has a high stagnation pressure. The portion of primary airflow45 within boundary layer 38 has a lower velocity and a correspondinglylower stagnation pressure. Accordingly, prior art propulsion system 20yields substantial radial distortion of the airflow reaching face 48.

Engine core 34 includes the primary powered components of engine 32. Forexample, engine core 34 may comprise a turbine, a combustor, and acompressor, among other components. For ease of illustrations, theseengine components are not illustrated. Engine fan bypass 36 is a regionthat is largely devoid of any powered components. Rather, this is aregion where the engine's fan pushes air around the outside of enginecore 34. As illustrated in FIG. 1, boundary layer 38 comprises asubstantial portion of the airflow that enters engine core 34.Accordingly, a substantial portion of the air that enters engine core 34has a stagnation pressure that is lower than the remainder of primaryairflow 45. In the illustrated embodiment, engine core 34 has a highersensitivity to changes in stagnation pressure than engine fan bypass 36.In other words, an increase in the stagnation pressure of the airflowentering engine core 34 would yield a relatively large increase inthrust whereas a corresponding increase in the stagnation pressure ofthe airflow entering engine fan bypass 36 would yield a smaller increasein thrust and, conversely a reduction in the stagnation pressureentering engine fan bypass 36 would result in a correspondingly smalldecrease in net thrust.

Accordingly, it would be desirable to direct the higher velocity air ofprimary airflow 45 towards the engine core and to direct the lowervelocity air of primary flow 45 (i.e., the air in boundary layer 38)away from the engine core.

FIG. 2 is a schematic cross sectional view of a propulsion system 50made in accordance with the teachings of the present disclosure. Withcontinuing reference to FIG. 1, propulsion system 50 is substantiallyidentical to prior art propulsion system 20 with the primary exceptionbeing that propulsion system 50 includes a plurality of vortexgenerators 52 disposed circumferentially about center body 22. It shouldbe understood that FIG. 2 presents a schematic representation and,accordingly, the depiction of each vortex generator has been simplifiedfor ease of illustration.

In the illustrated embodiment, each vortex generator of the plurality ofvortex generators has a height that exceeds the thickness of boundarylayer 38 at the locations where plurality of vortex generators 52encounter boundary layer 38. Consequently, each vortex generator ofplurality of vortex generators 52 will encounter the high velocity airof primary airflow 45 and will generate a plurality of vortices 54, withalternating sense of direction, that will propagate downstream indiffuser region 46. Depending on the height of plurality of vortexgenerators 54, plurality of vortices 54 may propagate through both thehigh velocity air and the slower moving air of boundary layer 38, orthey may propagate only through the high velocity air. If the VGs are ofsub-boundary layer height or are equal to boundary layer height, thenplurality of vortices 54 will propagate through both boundary layer 38and through the high velocity air outside of boundary layer 38. If thevortices are created outside of boundary layer 38, then they will remainthere.

As plurality of counter-rotating vortices 54 propagate, they willgenerate up-wash 56 and downwash 58 that will move the high velocity airof primary airflow 45 radially towards center body 22 and that will alsomove the low velocity air of primary airflow 45 (i.e., the air inboundary layer 38) radially away from center body 22. This radialmovement of the high and low velocity air of primary airflow 45 willcause the high velocity air and the low velocity air to mix andintermingle. This causes an exchange of energy between the differentportions of primary airflow 45 (higher velocity portion and lowervelocity portion) that yields an airflow having a more homogenousvelocity throughout primary airflow 45. This, in turn, reduces thedisparity of the stagnation pressure at different radial locationsthroughout primary airflow 45 and therefore lowers the overall radialdistortion of primary airflow 45 before it reaches face 48 of engine 32.Thus, as a result of the agitation of primary airflow 45 by plurality ofvortex generators 52, the stagnation pressure of the portion of primaryairflow 45 entering engine core 34 will be higher than it would havebeen had plurality of vortex generators 52 not been present.

FIG. 3 is a cross-sectional view taken along the line 3-3 of FIG. 2.Plurality of vortex generators 52 are arranged circumferentially aboutcenter body 22 in pairs. With continuing reference to FIGS. 1-2, and asdiscussed in greater detail below, each pair of vortex generators isoriented at a predetermined angle of attack with respect to primaryairflow 45. In some embodiments, each pair of vortex generators may bepositioned and aligned at substantially the same axial location alongcenter body 22. In other embodiments, the axial position of each pair ofvortex generators along center body 22 may be varied. In the illustratedembodiment, each vortex generator 52 of each pair of vortex generatorshas a rectangular configuration. Further, as illustrated by vortexgenerator 52 a, each vortex generator 52 has an aspect ratio of 2. Itshould be understood that this is merely exemplary in nature, in otherembodiments, plurality of vortex generators 52 may have any suitableshape, contour, aspect ratio, or configuration without departing fromthe teachings of the present disclosure.

With continuing reference to FIGS. 1-3, FIG. 4 is a fragmentaryschematic view illustrating a section of center body 22 viewed from theperspective of arrow 4 of FIG. 3. In this illustration, two pairs ofvortex generators 52 are depicted. Each vortex generator 52 is orientedat an angle of attack α. In various embodiments, angle of attack α mayvary from sixteen degrees to twenty four degrees. In other embodiments,angle of attack α may have any suitable magnitude. In the illustratedembodiment, each vortex generator 52 of each pair of vortex generatorsis canted at an equal and opposite angle. For example, vortex generator52 a is canted at an angle of attack α of positive 16 degrees and vortexgenerator 52 b is canted at an angle of attack α of negative 16 degrees.In other embodiments, the angles of attack α of each vortex generator 52of each pair may not be equal. In still other embodiments, the angles ofattack α of each vortex generator 52 of each pair may not be opposite(i.e., they may both be negative or they may both be positive).

In the illustrated embodiment, each vortex generator 52 of each pair ofvortex generators is arranged such that their respective downstreamsides (from the perspective of primary airflow 45) are closer togetherthan their respective upstream sides. Thus, as air from primary airflow45 passes between the two vortex generators 52 of each pair of vortexgenerators, the vortices generated by each pair of vortex generatorscreate an up-wash that will move slower moving air from boundary layer38 into the faster moving air of primary airflow 45. Similarly, as airfrom primary airflow 45 passes between each pair of vortex generators,vortices are generated that create a downwash that will move the fastermoving air from primary airflow 45 down towards center body 22. Thecombination of the up-wash and the downwash cooperate to break up theboundary layer, and mix, redistribute, and generally homogenize theflow. This creates a more consistent stagnation pressure throughoutprimary airflow 45.

FIG. 5 is a perspective view illustrating a portion of propulsion system50, viewed while looking in a downstream direction. In the illustratedembodiment, the vortex generators of each pair of vortex generators, andeach pair of vortex generators themselves are spaced apart at arelatively wide gap. FIG. 6 is a perspective view illustrating a portionof a propulsion system 50′, viewed while looking in a downstreamdirection. Propulsion system 50′ is substantially identical topropulsion system 50, the only difference being that the vortexgenerators of each pair of vortex generators and each pair of vortexgenerators themselves are spaced apart at a relatively narrow gap. Thespacing between vortex generators and the spacing between each pair ofvortex generators can be varied as needed for each application.

With continuing reference to FIGS. 1-6, FIG. 7 is an expanded schematicside view illustrating a portion of center body 22 with boundary layer38 flowing over the surface of center body 22. Also illustrated are twovortex generators, a vortex generator 52 x and a vortex generator 52 y,which may be utilized with propulsion system 50. Vortex generator 52 xand 52 y each have different heights. Vortex generator 52 x has a heightthat is approximately seventy five percent of the local thickness ofboundary layer 38 while vortex generator 52 y has a height that isapproximately one hundred and twenty five percent of the local thicknessof boundary layer 38. Vortex generators having a height of seventy fivepercent of the local thickness of the boundary layer will reach into theupper reaches of the boundary layer where the air is flowing at arelatively fast rate as compared with the remainder of the boundarylayer. Vortex generators having a height of anything over one hundredpercent of the local boundary layer thickness will protrude beyond theboundary layer and will encounter the fast moving portions of primaryairflow 45. Accordingly, FIG. 7 illustrates the lower and upper limits apreferred range of vortex generator heights, although it should beunderstood that the use of vortex generators having greater or lesserheights may also be employed without departing from the teachings of thepresent disclosure, so long as the vortex generators cause thegeneration of vortices that propagate at least partially outside of theboundary layer and cause the mixing and intermingling of high and lowvelocity air, as discussed above.

FIG. 8 is a schematic cross-sectional view of a propulsion system 60made in accordance with the teachings of the present disclosure. Withcontinuing reference to FIGS. 1-7, propulsion system 60 is substantiallyidentical to prior art propulsion system 20 with the primary exceptionbeing that propulsion system 60 includes a plurality of vortexgenerators 62 disposed circumferentially about an inner surface ofsplitter 30. It should be understood that FIG. 8 presents a schematicrepresentation and, accordingly, the depiction of each vortex generatorhas been simplified for ease of illustration. By positioning pluralityof vortex generators 62 on the internal surface of splitter 30, eachvortex generator 62 has greater access to the high velocity air ofprimary airflow 45 than do plurality of vortex generators 52 because thethickness of the boundary layer that forms on the internal surface ofsplitter 30 is relatively small as compared with boundary layer 38.Accordingly, if vortex generators are positioned on the internal surfaceof splitter 30 (or on the internal surface of nacelle 26 in propulsionsystems that do not employ a splitter), such vortex generators may besmaller than vortex generators positioned on center body 22 while stillachieving substantially the same benefit.

With continuing reference to FIGS. 1-7, in the embodiment illustrated inFIG. 8, each vortex generator of the plurality of vortex generators 62will encounter the high velocity air of primary airflow 45 and willgenerate a plurality of vortices 64 that will propagate through diffuserregion 46. As plurality of vortices 64 propagate, they will generateup-wash 56 and downwash 58 that will move the high velocity air ofprimary airflow 45 radially towards center body 22 and will also movethe low velocity air of primary airflow 45 (i.e., the air in boundarylayer 38) radially away from center body 22. As discussed above withrespect to vortex generators 52, this movement of the high and lowvelocity air of primary airflow 45 will cause the high and low velocityair to mix and intermingle. This, in turn, reduces the disparity of thestagnation pressure at different radial locations throughout primaryairflow 45 and lowers the overall radial distortion of primary airflow45 before it reaches face 48 of engine 32. Thus, as a result of theagitation of primary airflow 45 by plurality of vortex generators 62,the stagnation pressure of the portion of primary airflow 45 enteringengine core 34 will be higher than it would have been had plurality ofvortex generators 62 not been present.

FIG. 9 is a cross-sectional view taken along the line 9-9 of FIG. 8.Plurality of vortex generators 62 are arranged circumferentially inpairs about an inner surface 66 of splitter 30. With continuingreference to FIGS. 1-8, each pair of vortex generators is oriented at apredetermined angle of attack α with respect to primary airflow 45. Insome embodiments, each pair of vortex generators may be positioned andaligned at substantially the same axial location along inner surface 66.In other embodiments, the axial position of each pair of vortexgenerators may be varied. In the illustrated embodiment, each vortexgenerator 62 of each pair of vortex generators has a rectangularconfiguration. It should be understood that this is merely exemplary innature, in other embodiments, plurality of vortex generators 62 may haveany suitable shape, contour, aspect ratio or configuration withoutdeparting from the teachings of the present disclosure.

FIG. 10 is a schematic cross sectional view of a propulsion system 70made in accordance with the teachings of the present disclosure. Withcontinuing reference to FIGS. 1-9, propulsion system 70 is substantiallyidentical to prior art propulsion system 20 with the primary exceptionbeing that propulsion system 70 includes a plurality of vortexgenerators 52 disposed circumferentially about center body 22 andanother plurality of vortex generators 62 disposed circumferentiallyabout an inner surface of splitter 30. It should be understood that FIG.10 presents a schematic representation and, accordingly, the depictionof each vortex generator has been simplified for ease of illustration.By positioning plurality of vortex generators 52 circumferentially aboutcenter body 22 and also positioning plurality of vortex generators 62circumferentially on the internal surface of splitter 30, the two setsof vortex generators can cooperate to create a flow that is stronger,that has mutually reinforcing up-wash and downwash, and that leads tostronger radial gradients and more complete flow redistribution andmixing (both radially and circumferentially) than would be possibleusing only a single set of vortex generators at only one of the twolocations. This enhanced flow more effectively causes the mixing andintermingling of the fast moving portions and the slow moving portionsof primary airflow 45, and thereby yields a flow that is even morehomogenous when it reaches face 48 of engine 32 than would be producedby either set of vortex generators acting alone.

Each vortex generator of plurality of vortex generators 52 willencounter the high velocity air of primary airflow 45 and will generatea plurality of vortices 54 that will propagate downstream throughdiffuser region 46. Each vortex generator of the plurality of vortexgenerators 62 will also encounter the high velocity air of primaryairflow 45 and will generate a plurality of vortices 64 that willpropagate downstream through diffuser region 46. As plurality ofvortices 54 and 64 propagate, they will generate up-wash 56 and downwash58 that will move the high velocity air of primary airflow 45 radiallytowards center body 22 and will also move the low velocity air ofprimary airflow 45 (i.e., the air in boundary layer 38) radially awayfrom center body 22. This movement of the high and low velocity air ofprimary airflow 45 will cause the high and low velocity air to mix andintermingle. This, in turn, reduces the disparity of the stagnationpressure at different radial locations throughout primary airflow 45 andlowers the overall radial distortion of primary airflow 45 before itreaches face 48 of engine 32. Thus, as a result of the agitation ofprimary airflow 45 by plurality of vortex generators 62, the stagnationpressure of the portion of primary airflow 45 entering engine core 34will be higher than it would have been without plurality of vortexgenerators 62.

FIG. 11 is a perspective view illustrating a non-limiting embodiment ofa supersonic aircraft 80. Supersonic aircraft 80 includes a propulsionsystem 82. Propulsion system 82 includes an engine, a center bodydisposed upstream of the engine, a shroud partially surrounding theengine and the center body, the shroud configured to direct an airflowpassing over the center body towards the engine and a plurality ofvortex generators positioned upstream of the engine. The plurality ofvortex generators each have a height such that when supersonic aircraft80 is flown at a predetermined speed, the plurality of vortex generatorscreate a plurality of vortices that propagate at least partially outsideof a boundary layer that is formed proximate the center body. Theplurality of vortices cause a high-velocity portion of the airflow tomove radially towards the center body and a low-velocity portion of theairflow to move radially away from the center body prior to the airflowreaching the AIP.

With continuing reference to FIGS. 1-10, in some non-limitingembodiments, propulsion system 82 may be substantially identical topropulsion system 50. In other non-limiting embodiments, propulsionsystem 82 may be substantially identical to propulsion system 60. Instill other non-limiting embodiments, propulsion system 82 may besubstantially identical to propulsion system 70. While supersonicaircraft 80 is depicted as a fixed wing aircraft, it should beunderstood that aircraft having any suitable configuration may also beemployed without departing from the teachings of the present disclosure.Similarly, although supersonic aircraft 80 has been illustrated with twonacelles 82 attached to a vertical stabilizer, it should be understoodthat any suitable number of nacelles 80 may be employed and further,that they may be housed at any suitable location on the aircraft.

While at least one exemplary embodiment has been presented in theforegoing detailed description of the disclosure, it should beappreciated that a vast number of variations exist. It should also beappreciated that the exemplary embodiment or exemplary embodiments areonly examples, and are not intended to limit the scope, applicability,or configuration of the invention in any way. Rather, the foregoingdetailed description will provide those skilled in the art with aconvenient road map for implementing an exemplary embodiment of theinvention. It should be understood that various changes may be made inthe function and arrangement of elements described in an exemplaryembodiment without departing from the scope of the disclosure as setforth in the appended claims.

What is claimed is:
 1. An arrangement for use with a propulsion systemfor a supersonic aircraft, the arrangement comprising: a center bodyconfigured for coupling to an inlet and to support a boundary layerformed when the supersonic aircraft is flown at a predetermined altitudeand a predetermined speed greater than Mach 1; a first vortex generatordisposed on the center body, the first vortex generator extending afirst height above the center body; and a second vortex generatordisposed on the center body, said second vortex generator extending asecond height above the center body, the second height being greaterthan the first height, wherein said first height and said second heightare greater than approximately seventy-five percent of a thickness ofthe boundary layer proximate a location of the first vortex generatorand the second vortex generator, respectively, when the aircraft ifflown at the predetermined altitude and the predetermined speed.
 2. Thearrangement of claim 1, wherein one of the first height and the secondheight are greater than 100% of the thickness of the boundary layer. 3.The arrangement of claim 1, wherein one of the first height and thesecond height are greater than one hundred and twenty percent of thethickness of the boundary layer.
 4. The arrangement of claim 1, whereinthe first height and the second height are greater than one hundred andtwenty-five percent of the thickness of the boundary layer.
 5. Thearrangement of claim 1, wherein the first vortex generator and thesecond vortex generator are disposed circumferentially about the centerbody.
 6. The arrangement of claim 1, wherein the second vortex generatorforms a first generator pair with the first vortex generator.
 7. Thearrangement of claim 1, wherein the first vortex generator has asubstantially rectangular shape.
 8. The arrangement of claim 1, whereinthe first vortex generator has an aspect ratio of approximately
 2. 9. Apropulsion system for a supersonic aircraft, the propulsion systemcomprising: a plurality of vortex generators positioned upstream of anengine of the propulsion system, each vortex generator of the pluralityof vortex generators having a height such that when the supersonicaircraft is flown at a predetermined altitude and at a predeterminedcruise speed, the predetermined cruise speed being greater than Mach 1,the plurality of vortex generators create a plurality of vortices thatpropagate at least partially outside of a boundary layer formedproximate a surface of a supersonic inlet, the plurality of vorticescausing a high-velocity portion of the airflow to move towards theengine core and a low-velocity portion of the airflow to move towardsthe engine bypass prior to the airflow reaching a face of the engine,the plurality of vortex generators being disposed aft of a terminalshock formed when the supersonic aircraft is flown at the predeterminedcruise speed, and each vortex generator of the plurality of vortexgenerators having a height greater than a thickness of the boundarylayer at the location of each vortex generator of the plurality ofvortex generators.
 10. The propulsion system of claim 9, wherein theheight of each vortex generator of the plurality of vortex generatorsdoes not exceed one hundred and twenty-five percent of the thickness ofthe boundary layer at the location of each vortex generator of theplurality of vortex generators.
 11. The propulsion system of claim 9,wherein the plurality of vortex generators are arrangedcircumferentially about the surface of the supersonic inlet.
 12. Thepropulsion system of claim 9, wherein the plurality of vortex generatorsare arranged circumferentially about an inner surface of a shroud of thepropulsion system.
 13. The propulsion system of claim 9, wherein theplurality of vortex generators are arranged circumferentially about boththe surface of the supersonic inlet and an inner surface of a shroud ofthe propulsion system.
 14. The propulsion system of claim 9, wherein theplurality of vortex generators are arranged in pairs of vortexgenerators.
 15. The propulsion system of claim 14, wherein each vortexgenerator of each pair of vortex generators is oriented with respect tothe airflow so as to have an angle of attack.
 16. The propulsion systemof claim 15, wherein the angle of attack is between sixteen degrees andtwenty-four degrees.
 17. The propulsion system of claim 15, wherein eachvortex generator of each pair of vortex generators is oriented atopposite angles of attack to each other.
 18. The propulsion system ofclaim 9, wherein the plurality of vortex generators are arranged suchthat the vortices generate both an up-wash and a downwash.
 19. Thepropulsion system of claim 9, wherein the predetermined altitudecomprises a predetermined cruise altitude.
 20. A propulsion system for asupersonic aircraft, the propulsion system comprising: a plurality ofvortex generators positioned upstream of an engine of the propulsionsystem, each vortex generator of the plurality of vortex generatorshaving a height such that when the supersonic aircraft is flown at apredetermined altitude and at a predetermined cruise speed, thepredetermined cruise speed being greater than Mach 1, the plurality ofvortex generators create a plurality of vortices that propagate at leastpartially outside of a boundary layer formed proximate a surface of asupersonic inlet, the plurality of vortices causing a high-velocityportion of the airflow to move towards the surface and a low-velocityportion of the airflow to move away from the surface prior to theairflow reaching a face of the engine, the plurality of vortexgenerators being disposed aft of a terminal shock formed when thesupersonic aircraft is flown at the predetermined cruise speed, and eachvortex generator of the plurality of vortex generators having a heightgreater than a thickness of the boundary layer at the location of eachvortex generator of the plurality of vortex generators.